This invention relates to a combustor for a gas turbine engine, and particularly, although not exclusively, for a gas turbine aeroengine.
The thrust generated by a gas turbine engine is modulated by varying the flow of fuel to the combustor or combustors. Efficient combustion requires the air/fuel ratio of the air/fuel mixture to be maintained within close limits. Efficient combustion is desirable both because it minimises undesirable emissions, such as NOx and CO emissions, and because it improves specific fuel consumption (SFC).
It is known to provide air flow restricting mechanisms at the combustor inlet so as to vary the quantity of air available for mixture with the fuel. However, such mechanisms can be complex and are not suited to operation in the hostile environment which exists at the combustor inlet.
US 2005/095542 discloses a gas turbine engine combustor having a variable-geometry air inlet for supplying air to a pre-mixing zone of the combustor. US 2005/095542 also discloses dilution ports in a liner of the combustor, which ports have adjustable flow areas controlled by valves.
Since the dilution ports are at separate locations around the combustor liner, the air flow admitted through them to the combustor is not distributed evenly, and consequently temperature differentials around the axis of the combustor can arise owing to the introduction of low temperature air at discrete positions around the axis of the combustor.